Thermodynamic reaction drive

ABSTRACT

A thermodynamic reaction drive or rocket engine which employs liquefiable solid propellant. An alkali metal, which is used as the propellant, is stored in the fuel tank of the rocket engine in the solid state and then, whenever engine operation is desired, liquefied by heating to the necessary temperature. The liquid propellant is fed from the fuel tank through a De Laval nozzle into the mixing or combustion chamber where it is combined with an oxidizer. The propellant, which can be forced through the nozzle primarily by means of its own vapor pressure, is changed over at the nozzle into a two-phase flow. The oxidizer, which is simultaneously introduced into the mixing chamber at a higher pressure than the pressure in the chamber, effects the combustion of the propellant so that the internal energy of the products of combustion may be converted to kinetic energy in an appropriate thrust nozzle.

United States Patent [72] lnventors Reinhart Radebold;

Hermann Lang, both 01 Berlin, Germany [21] Appl. No. 835,279 [22] FiledApr. 9, 1969 Division of Ser. No. 717,287. Mar. 29, 1968. [45] PatentedJuly 27, 1971 [73] Assignee Licentia Pntent-Verwaltungs G.m.b.H.

Frankfurt am Main, Germany {32] Priority Apr. 1, 1967,.luly 11,1967 [33]Germany [31] L56 149andL56963 [54] THERMODYNAMIC REACTION DRIVE 10Claims, 4 Drawing Figs.

[52] US. Cl 60/258, 60/39.46, 60/39.48, 60/259, 60/267, 60/270 [51] Int.Cl F02k 9/02, F02k 7/10 [50] Field ofSearch 150/3948, 259, 258, 264,39.46, 257, 267, 270

[56] References Cited UNITED STATES PATENTS 2,408,1 ll 9/1946 Truax60/39.48 3,102,388 9/1963 Abild.... 60/259 3,107,485 10/1963 Toulmin60/260 3,116,603 l/ 1964 Hausmann 60/224 3,149,460 9/1964 Rocca 60/39.46

3,203,171 8/1965 Burke 60/21 1 3,320,742 5/1967 Truax 60/39.48

3,368,354 2/1968 Adelman 60/39.72

FOREIGN PATENTS 276,91 1 8/1930 ltaly 60/260 1,004,298 11/1951 France60/258 Primary Examiner-Douglas Hart Attorney-Spencer & Kaye ABSTRACT: Athermodynamic reaction drive or rocket en gine which employs liquefiablesolid propellant. An alkali metal, which is used as the propellant, isstored in the fuel tank of the rocket engine in the solid state andthen, whenever engine operation is desired, liquefied by heating to thenecessary temperature. The liquid propellant is fed from the fuel tankthrough a De Laval nozzle into the mixing or combustion chamber where itis combined with an oxidizer. The propellant, which can be forcedthrough the nozzle primarily by means of its own vapor pressure, ischanged over at the nozzle into a two-phase flow. The oxidizer, which issimultaneously introduced into the mixing chamber at a higher pressurethan the pressure in the chamber, effects the combustion of thepropellant so that the internal energy of the products of combustion maybe converted to kinetic energy in an appropriate thrust nozzle.

PATENTEDJULN WI 3 5915 022 sum 1 or 2 Fig] Reinyqrt Rondabold HermannLung '32, Q wu/z 4/140434 THERMODYNAMIC REACTION DRIVE CROSS-REFERENCETO RELATED APPLICATION This application is a division of our copendingapplication Ser. No. 717,287, filed March 29th 1968, entitled THER-MODYNAMIC REACTION DRIVE.

BACKGROUND OF THE INVENTION The present invention relates to athermodynamic reaction drive or rocket engine.

The rocket engines known in the art, which are used, for example, inspacecraft, operate with either solid or liquid propellants. Solidpropellants have the disadvantage that, once the rocket engine isignited, its thrust can not be controlled and must continue until all ofthe propellant has been oxidized. A rocket engine utilizing liquidpropellant can, in fact, be controlled by apportioning, e.g., by meansof pumps, the amount of propellant which is permitted to react. Theutilization of liquid propellant such as liquid hydrogen and oxygen,however, requires that certain safety measures be taken to avoidexplosions. A further disadvantage of the liquid propellant rocketengine is that, when installed in a spacecraft, it requires an extensiveground installation and, because the liquid propellant can notcontinuously be maintained in the engines fuel tanks, it requires anelaborate starting procedure.

It is also known in the art to employ as a propellant a mixture of ametal powder and a liquid reaction partner. To guarantee a homogeneousmix this propellant mixture is embedded in a gel which is normally solidbut which becomes fluid or forms an emulsion when it is pumped to beinjected into the combustion chamber. However, even this type of rocketengine requires considerable technical effort in its design andconstruction.

SUMMARY OF THE INVENTION An object of the present invention, therefore,is to provide a thermodynamic reaction drive, or, in particular, arocket engine for spacecraft, which does not admit of the disadvantagesdescribed above and which may be realized by apparatus which isrelatively easy to construct.

This, as well as other objects which will become apparent in thediscussion that follows, is achieved, according to the presentinvention, by employing an alkali metal as the propellant. Thispropellant is stored in the rocket engine fuel tank in the solid stateand then, when engine operation is desired, liquefied by heating to thenecessary temperature. The liquid propellant is fed from the fuel tankthrough a De Laval nozzle into the mixing or combustion chamber where itis combined with an oxidizer. The propellant, which can be forcedthrough the nozzle primarily by means of its own vapor pressure, ischanged over at the nozzle into a two-phase flow. The oxidizer, which issimultaneously introduced into the mixing chamber at a higher pressurethan the pressure in the chamber, effects the combustion of thepropellant so that the internal energy of the products of combustion maybe converted to kinetic energy in an appropriate thrust nozzle. BRIEFDESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic diagram of rocketengine apparatus according to one preferred embodiment of the presentinvention.

FIG. 2 is a schematic diagram of rocket engine apparatus according toanother preferred embodiment of the present invention.

FIG. 3 is a schematic diagram of the propellant injecting portion ofrocket engine apparatus according to another preferred embodiment ofthepresent invention.

FIG. 4 is a schematic diagram of rocket engine apparatus according toanother preferred embodiment of the present invention.

2 DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to the drawings,FIG. 1 illustrates one embodiment of the rocket engine apparatusaccording to the present invention. The fuel tank l in FIG. 1 isconnected via a De Laval or converging-diverging nozzle 2 with a mixingor, in this case, combustion chamber 3. The De Laval nozzle projectsinto the combustion chamber. A thrust nozzle 4, which is likewiseconstructed as a De Laval nonle, is attached to the combustion chamber3. The fuel or propellant 6 which may, for example, be potassium, isintroduced into the fuel tank in the liquid state and allowed tosolidify. The De Laval nozzle 2 is initially closed in the manner knownin the art by a diaphragm 5 held in a groove in the nozzle periphery.

If the potassium propellant in the fuel tank is brought to a temperatureof about 900 C. by means of heating apparatus 1" it will liquefy and,with the engine apparatus arranged in the vertical position, will form avapor space 7 at the top of the tank that has a pressure ofapproximately 20 atmospheres.

The notch in which the diaphragm 5 is mounted is so dimensioned that thediaphragm will automatically blow off at this pressure. The liquidpotassium will then enter the mixing chamber in finely distributeddroplets. A thermodynamic atomization of the potassium will take placeas a result of its own vapor. If it be assumed that the expansion in theDe Laval nozzle 2 is adiabatic, it follows that small bubbles ofsuperheated vapor will be produced as a result of the pressure drop,which bubbles will distribute the kinetic'energy of the fluid as theyexpand forming an almost homogeneous two-phase flow. This thermodynamicatomization can thus provide about 20 percent of the total thrust of therocket engine; that is, sufficient thrust to cause the engine to hover.

The oxidizing agent or oxidizer 9 in the container 8 is led throughregulator valves 11, a collecting ring 12 and the injector nozzles 13into the combustion chamber. The nozzles 13 can be arranged in a ring,for example, in the manner well known in the art. If necessary, theoxidizer may be preheated prior to injection so that its pressure willlie above the ambient pressure in the combustion chamber.

The internal energy of the products of combustion produced by. theoxidation will be converted into kinetic energy in the thrust nozzle 4.

A halogen is preferably employed as oxidizing agent in this rocketengine. By using a halogen, fluorine or chlorine, in particular, it ispossible to obtain the necessary vapor pressure by preheating, asmentioned above, making it unnecessary to employ injection pumps.

It is possible to use any substance as an oxidizer which will react withthe alkali metal used as fuel. Some of the other oxidizers which willspontaneously react with an alkali metal are oxygen, halogens andalcohols.

The oxidation of the propellant can also be achieved using environmentaloxygen during flights in the region of the earths atmosphere. Thenecessary pressure can, in this case, be developed by means of anair-catclhing diffusor Ml as illustrated in FIG. 2. The starting speedof the rocket engine which is required to render the diffusor operativecan be produced, for example, by an initial rocket stage or by someother wellknown means for accelerating the apparatus on the ground.

It is possible also to provide the engine shown in FIG. 2 with theadditional oxidizing apparatus of the type illustrated in FIG. I so thathalogen can again be used as the oxidizer outside the atmosphere. Theutilization of a diffusor makes it possible to exclude an otherwisenecessary quantity of oxidizing agent and thus considerably reduce thetakeoff weight of the spacecraft.

The utilization of an alkaline metal as the propellant, according to thepresent invention, permits the thrust of the rocket engine to beswitched off or controlled in an advantageous manner. FIG. 3 illustratesone embodiment of the propellant supply system by which this may beaccomplished. In this embodiment the propellant 6 is supplied to the DeLaval nozzle 2 from the fuel tank ll through lines l5 and regulatingvalves 16. This type of propellant supply system does not reduce theefficiency of the rocket engine. The combustion chamber and the oxidizersupply system have not been included in FIG. 3 in the interestofclarity.

In a further advantageous embodiment of the present invention alkalinemetal consists of a homogeneous mixture of two alkali elements, one ofwhich, the principal constituent, has a higher specific heat of reactionand a higher boiling temperature than the other, secondary constituent.The secondary constituent thus effects the thermodynamic atomization ofthe principal constituent, whereas the heat of reaction of the latterdelivers the main thrust of the rocket engine. The temperature of thefuel tank during operation is so chosen, in dependence upon the pressurein the mixing chamber, that the secondary constituent boils when itexpands in the De Laval nozzle and causes the principal constituent toatomize.

A mixture of potassium and lithium provides especially good results as atwo-constituent propellant. A l-2() per cent share of potassium in thetotal mixture is sufficient here to uniformly and finally atomizc thelithium.

According to a still further embodiment of the present invention thepropellant may also be inserted into the mixing chamber with the aid ofa separate gas. This measure allows the operating temperature of thepropellant to be reduced. If the gas, which, for example, may be argon,is employed at a pressure of about 20 atmospheres the normally necessarypropellants temperature of 900 C. may be reduced to approximately 600 C.This gas may be introduced into the fuel tank I through an inlet 18 atthe top of the tank, as shown in FIG. 1, after the propellant has beenliquefied.

It is possible to increase the thrust of the rocket engine if theexpansion in the thrust nozzle is made essentially isothermal. This canbe achieved by providing a number of oxidizer inlet openings 19 alongthe longitudinal surface of the combustion chamber so that thecombustion process is continued up to the exit opening of the thrustnozzle, as shown in FIG. 1.

The heat generated by the rays of the sun on a launching pad may, forexample, be sufficient to liquefy the propellant. In regions where asufficient amount of solar energy is not normally available it ispossible, according to another embodiment of the present invention, tointroduce a halogen oxidizer into the solid propellant to liquefy it andraise it to a higher temperature. It is also practical, in addition, tolead the liquefied propellant through a heat exchanger formed by theexternal surface of the combustion chamber to further increase itstemperature before it enters the De Laval nozzle.

A rocket engine constructed according to this embodiment isschematically illustrated in FIG. 4.

Assuming that the propellant 6 is present in the fuel tank 1 in theliquid state, this propellant can pass through the valve I6 into a heatexchanger 17 which consists of an external wall, a guide plate, and theexternal surface of the combustion chamber 3. When the propellantemerges from the heat exchanger it has attained the temperature of 800to lO00 C., which is necessary for the operation of the engine.

To bring the rocket engine into the state of operational readiness theinitially solid propellant 6 is liquefied by a halogen which serves asan oxidizer. The propellant may be heated in this way, for example, to amean temperature of about 200 to 300 C. The halogen may, for example, bein troduced through the holes in the fuel tank and in the solid alkalimetal.

In the embodiment shown in FIG. 4 the internal wall of the fuel tank isprovided with distributing rings 20. If the propellant be introducedinto the fuel tank in the liquid state through an opening, not shown, atthe lower end of the tank, an overpressure formed by the rings willcreate cavities inside these rings in the region of the tank wall.Closable side openings in the container wall next to these cavitiespermit the introduction of the halogen, preferably fluorine, when theengine is to be brought into operational readiness. The resultantliberated heat of oxidation liquifies the propellant and brings it up toa mean temperature. It is advantageous if the same halogen which is usedto liquefy the propellant be also used for the reaction in thecombustion chamber.

The embodiment illustrated in FIG. 4 and described above is additionallyadvantageous in that it permits a considerable reduction in thethickness of the wall of the fuel tank and therefore in the weight ofthe rocket engine. This embodiment simultaneously affords anadvantageous cooling of the mixing chamber.

It will be understood that the above description of the presentinvention is susceptible to various modifications, changes, andadaptations, and the same are intended to be comprehended within themeaning and range of equivalents of the appended claims.

We claim:

1. Reaction drive apparatus comprising, in combination:

a. a fuel tank:

b. a mass of alkali metal propellant contained in said tank;

c. a mixing chamber;

d. first De Laval nozzle means connected to said fuel tank fordischarging fuel into said mixing chamber;

e. heating means for heating said propellant in said tank to atemperature sufficient for causing said propellant to be dischargedthrough said nozzle in a two-phase liquid gas flow; and

f. means for injecting an oxidizer for said propellant into said mixingchamber.

2. The apparatus defined in claim I, wherein said mixing chamber is acombustion chamber.

3. The apparatus defined in claim 2, further comprising second De Lavalnozzle means connected to said combustion chamber for dischargingproducts of combustion from said combustion chamber.

4. The apparatus defined in claim 3, wherein said first De Laval nozzlemeans is arranged to discharge fuel in the same direction as said secondDe Laval nozzle means discharges products of combustion.

5. The apparatus defined in claim 1, further comprising control valvemeans arranged between said fuel tank and said first De Laval nozzlemeans for regulating the amount of fuel passing from said fuel tank tosaid first De Laval nozzle means.

6 The apparatus defined in claim 1, wherein said means for injecting anoxidizer into said mixing chamber includes diffusor means for catchingatmospheric air and injection said air into said mixing chamber.

7. The apparatus defined in claim 3, wherein said means for injecting anoxidizer into said mixing chamber includes injection nozzle meansarranged to inject oxidizer at least in the region where said second DeLaval nozzle is connected to said mixing chamber.

8. The apparatus defined in claim 7, wherein said injection nozzle meansis arranged to inject oxidizer along the entire length of said mixingchamber.

9. The apparatus defined in claim 2, further comprising heat exchangermeans connected to said fuel tank, said first De Laval nozzle and tosaid combustion chamber for preheating said fuel after it leaves saidfuel tank and prior to its discharge through said first De Laval nozzle,said heat exchanger using heat generated in said combustion chamber.

10. An arrangement as defined in claim 1, wherein said propellant isnormally present in its solid state.

1. Reaction drive apparatus comprising, in combination: a. a fuel tank:b. a mass of alkali metal propellant contained in said tank; c. a mixingchamber; d. first De Laval nozzle means connected to said fuel tank fordischarging fuel into said mixing chamber; e. heating means for heatingsaid propellant in said tank to a temperature sufficient for causingsaid propellant to be discharged through said nozzle in a two-phaseliquid gas flow; and f. means for injecting an oxidizer for saidpropellant into said mixing chamber.
 2. The apparatus defined in claim1, wherein said mixing chamber is a combustion chamber.
 3. The apparatusdefined in claim 2, further comprising second De Laval nozzle meansconnected to said combustion chamber for discharging products ofcombustion from said combustion chamber.
 4. The apparatus defined inclaim 3, wherein said first De Laval nozzle means is arranged todischarge fuel in the same direction as said second De Laval nozzlemeans discharges products of combustion.
 5. The apparatus defined inclaim 1, further comprising control valve means arranged between saidfuel tank and said first De Laval nozzle means for regulating the amountof fuel passing from said fuel tank to said first De Laval nozzle means.6 The apparatus defineD in claim 1, wherein said means for injecting anoxidizer into said mixing chamber includes diffusor means for catchingatmospheric air and injection said air into said mixing chamber.
 7. Theapparatus defined in claim 3, wherein said means for injecting anoxidizer into said mixing chamber includes injection nozzle meansarranged to inject oxidizer at least in the region where said second DeLaval nozzle is connected to said mixing chamber.
 8. The apparatusdefined in claim 7, wherein said injection nozzle means is arranged toinject oxidizer along the entire length of said mixing chamber.
 9. Theapparatus defined in claim 2, further comprising heat exchanger meansconnected to said fuel tank, said first De Laval nozzle and to saidcombustion chamber for preheating said fuel after it leaves said fueltank and prior to its discharge through said first De Laval nozzle, saidheat exchanger using heat generated in said combustion chamber.
 10. Anarrangement as defined in claim 1, wherein said propellant is normallypresent in its solid state.